CN117999405A

CN117999405A - Three-stream turbine structure

Info

Publication number: CN117999405A
Application number: CN202180100289.2A
Authority: CN
Inventors: R·H·P·普林西瓦莱
Original assignee: Safran Aero Boosters SA; General Electric Co
Current assignee: Safran Aero Boosters SA; General Electric Co
Priority date: 2021-06-18
Filing date: 2021-06-18
Publication date: 2024-05-07
Also published as: WO2022262998A1; EP4355989A1

Abstract

The invention relates to a turbomachine (2) of the type having an unducted fan (4), comprising: a separation nozzle (10) that separates the air flow (F) into a main flow (F1) and a secondary flow (F2); a compressor (14) that compresses a main flow (F1); an air/oil heat exchanger (24); wherein the exchanger (24) is arranged in a channel (26) traversed by the third flow (F3), the third flow (F3) being taken from the secondary flow (F2) upstream of the exchanger (24), and the third flow encountering at least one annular row of rotor blades (20, 22) of the compressor (14) downstream of the exchanger (24).

Description

Three-stream turbine structure

Technical Field

The present invention relates to turbine design, and more particularly to an unducted fan type turbine. The present invention relates to an arrangement of heat exchangers for cooling turbine oil.

Background

In ducted turbines (turbojet engines), it is known to have one or more heat exchangers in the secondary flow (i.e. downstream of the fan).

The absence of a cowling around the secondary flow in a turboprop engine makes this arrangement impractical, and the heat exchanger is therefore supplied with cold air by a fan external to the engine. This solution is neither compact nor lightweight and also reduces the efficiency of the motor, since a part of the energy of the motor is necessary to drive the additional fan.

Application WO 2020/084271 A1 describes a so-called fast propeller or "open rotor" motor (for example of the "counter-rotating open rotor" type, or "unducted single fan (unducted SINGLE FAN, USF)" type). In this type of configuration, the air flow velocity is too low to cool sufficiently at low aircraft speeds and too high at cruising speeds, resulting in energy losses due to the drag of the exchanger. Thus, the known solutions for turbojet engines are not suitable for this type of engine and require very large external fans to ensure significant heat exchange requirements.

In short, the known technical solutions employ air directed towards the exchanger, either air that is too hot to cool the oil effectively or air that is too high in speed to ignore aerodynamic or thrust losses.

In addition to this, there are limitations associated with the vulnerability of the heat exchanger, which cannot be arranged directly upstream of the turbine, in particular because of the risk of collision of the heat exchanger with foreign bodies that may penetrate into the turbine.

Disclosure of Invention

Technical problem

The present invention aims to address the drawbacks of the design/manufacture of turbines in the prior art. In particular, the present invention aims to propose a structure that enables effective cooling in a limited space without impeding the efficiency of the turbine.

Technical proposal

The invention relates to a turbine of the unducted fan type, comprising: a separation nozzle that separates the air flow into a main flow and a secondary flow; a compressor that compresses a main stream; an air/oil heat exchanger; wherein the exchanger is arranged in a passage traversed by the third flow, the third flow is taken from the secondary flow upstream of the exchanger, and the third flow encounters at least one annular row of rotor blades of the compressor downstream of the exchanger.

The turbine may take the form of a turboprop or an open rotor (e.g. "counter-rotating open rotor, CROR", or "unducted single fan, USF") turbine.

The air flow is generated by a propeller and/or by the movement of an aircraft in which the turbine is mounted.

The unducted fan may be located upstream or downstream of the separation nozzle.

The compressor may be a low pressure compressor or a booster compressor.

The main flow is compressed by at least some of the compressor blades over at least a portion of its stroke. The secondary flow "sees" no compressor. The secondary flow is virtually infinite in size as there is no shroud around the secondary flow (radially outward). The third stream is different from the secondary stream. The third stream may converge with and/or diverge from the main stream.

At least one of the rotor blade rows of the compressor is located downstream of the heat exchanger, in other words, this means that the compressor takes air from the secondary flow and drives the air through the heat exchanger.

This so-called downstream portion of the compressor is located in an annular region where the main flow and the third flow converge.

This arrangement enables the exchanger to be supplied with air sufficiently cold and slow enough to ensure oil cooling efficiency and limited aerodynamic losses due to the presence of the exchanger.

According to an advantageous embodiment of the invention, the compressor comprises an upstream portion with at least one annular row of blades and a downstream portion with at least one annular row of blades, and only the downstream portion of the compressor is crossed by air from the third flow coming from the exchanger. This keeps the main stream, which increases rapidly in pressure and velocity, separate from the third stream. The heat exchanger "sees" no upstream portion of the compressor, thus maintaining the cooling capacity of the heat exchanger.

According to an advantageous embodiment of the invention, the radial height of the blades of the downstream portion is substantially greater than the radial height of the blades of the upstream portion, preferably the blades of the downstream portion of the compressor are 1.5 to 4 times radially higher than the blades of the upstream portion. In this way, the third flow does not suffer from a significant velocity gradient at the exchanger outlet, which would lead to aerodynamic turbulence, resulting in a loss of cooling efficiency and engine output.

According to an advantageous embodiment of the invention, the downstream portion of the compressor comprises a single annular row of rotor blades, the air encountering the downstream portion of the compressor downstream of the exchanger. This makes the whole unit more compact. The blades may be arranged in the form of bladed wheels. Further, the downstream portion may include one or two rows of stator vanes. The number of blades in the downstream portion may be set according to the available space and/or the desired compression ratio to both draw in the third stream at the desired speed and not to impede compression of the main stream, which is a key parameter for the efficiency of the turbine.

According to an advantageous embodiment of the invention, the air flow leaving the exchanger passes through all annular rows of compressor blades.

According to an advantageous embodiment of the invention, the opening is such that a part of the secondary flow is sucked in to form the third flow, the opening being non-ingestible (non-e copante). Thus, in some way, the opening may be defined by an upstream edge and a downstream edge, the upstream edge and the downstream edge having the same radial position. Alternatively, the upstream edge is farther from the turbine axis than the downstream edge, for example 1.1 times more. Alternatively, a tangent to the air guiding surface upstream of the opening describes a direction approaching the axis, as seen in axial section, and the opening is comprised between the axis and the tangent. In any case, this prevents the opening creating the third flow from forcing the exosomes into the exchanger in a scooping manner.

According to an advantageous embodiment of the invention, the channels and/or exchangers extend circumferentially over 360 ° around the axis of the turbine. Alternatively, the exchanger and/or the channels do not extend over 360 °. The opening to the secondary flow may be partial (i.e. not 360 °) and/or the mouth of the channel on the compressor blade may be partial (i.e. not 360 °) regardless of the angle of the exchanger about the axis.

According to an advantageous embodiment of the invention, the propeller is arranged upstream of the nozzle and/or the vanes straightening the secondary flow axially overlap with the downstream part of the compressor. Alternatively, the propeller may be arranged downstream of the nozzle. Alternatively, two propellers with opposite rotational directions are arranged at an axial position downstream of the exchanger.

According to an advantageous embodiment of the invention, a bypass is arranged downstream of the downstream portion of the compressor to divert the portion of the flow leaving the compressor towards the secondary flow. Thus, the bypass is located immediately downstream of the region of convergence of the main stream and the third stream (when the main stream and the third stream converge). This accelerates the flow to generate thrust and compensates for the air volume captured in the third flow upstream of the exchanger. This improves motor efficiency. The bifurcation of the bypass toward the secondary flow also enables any extraneous matter that may be ingested upstream of the primary and third flows to be discharged to prevent the extraneous matter from entering the high pressure compressor or combustor downstream of the primary flow.

According to an advantageous embodiment of the invention means are provided for isolating the main flow from the third flow, which flows through the downstream part of the compressor into the passage and then into the bypass. In this way, the third flow is completely isolated from the main flow, limiting disturbances downstream of the exchanger. However, the two flows are not completely independent in dynamics, since both flows "observe" the corresponding portions of the same row of rotor blades downstream of the exchanger.

According to an advantageous embodiment of the invention, at least the last annular row of blades of the downstream portion has an intermediate circumferential ring. The last row may consist of rotating blades or stationary blades. The axial ring directs the flow to a bypass or high pressure compressor. In combination with the previous section, the ring may also help isolate the third stream from the main stream.

The invention also relates to a method for cooling oil of a turbine with an unducted fan, the method comprising: dividing the air flow into a main flow and a secondary flow by a separating nozzle, the main flow being compressed by one or more compressors of the turbine, the secondary flow being located outside the one or more compressors; and forming a third flow through the air/oil exchanger by at least one annular row of rotor blades of the compressor, the third flow being taken from the secondary flow.

Finally, the invention relates to a method of using a turbine according to any of the above embodiments, the method comprising a propeller rotation step during which the mach number of the primary and secondary flows is 0.5 and the mach number of the third flow is well below 0.3.

Provides the advantages of

A particularly advantageous advantage of the present invention is that it enables cool air to circulate through the exchanger at a suitable rate to ensure effective cooling without compromising engine efficiency or requiring additional cumbersome equipment.

Good cooling efficiency means that the exchanger used is smaller and therefore smaller, lighter in weight and lower in cost.

Drawings

Fig. 1 shows a first embodiment of the invention;

Fig. 2 shows a second embodiment of the invention;

Fig. 3 shows a third embodiment of the invention;

Fig. 4 shows a fourth embodiment of the invention.

Detailed Description

In the following description, the terms "inner" and "outer" refer to positioning relative to the rotational axis of the turbine. The axial direction corresponds to a direction along the rotational axis of the turbine. The radial direction is perpendicular to the axis of rotation. Upstream and downstream refer to the direction of flow in the turbine.

The figures schematically illustrate elements and are not necessarily drawn to scale. In particular, some dimensions are exaggerated to make the drawing easier to read.

Fig. 1 shows a turbine 2 according to a first variant. The propeller 4 attached to the hub 6 rotates about an axis 8.

The turbine 2 moves in an air flow F, the movement of which with respect to the turbine 2 is caused by the rotation of the propeller 4 and the forward movement of the aircraft in which the turbine 2 is mounted.

In a variant not shown (which may be similar to one of the examples described in document WO 2020/084271 A1), the propeller 4 is arranged in the downstream portion of the turbine 2 and may optionally be supplemented by a second propeller having an opposite direction of rotation.

The air flow F is split into a main flow F1 and a secondary flow F2 at the separation nozzle 10.

The primary flow F1 encounters the rectifier inlet vanes (aube d' entr e e de redresseur, IGV) 11 and enters the flow passage 12, while the secondary flow F2 remains radially outward of any shroud. The housing 13 defines the outside of the flow channel 12. The stationary housing and the rotating hub portion define a flow passage 12 internally. The structural arms (posts) (see 13.1 in fig. 2) also pass through the flow channel 12 and are subjected to the forces of the housing 13.

The compressor 14 is designed to compress the main flow F1. To achieve this, the compressor 14 is equipped with alternating rotor blades 16, 18, 20, 22 and stator blades 17, 19, 21, which are arranged in annular rows about the axis 8.

Fig. 1 shows only the upstream portion of the turbine. Downstream of the compressor 14, the main flow F1 continues to a second compressor, a combustion chamber and one or more turbines (not shown). Rotation of the one or more turbines drives rotation of the hub 6, the propeller 4, and the rotor blades 16, 18, 20, 22.

The rotating elements are supported by bearings and the turbine may include a gearbox between the different rotating elements. The bearings and gearbox of the turbine 2 are lubricated by oil that must be kept within a given operating temperature range. Thus, a heat exchanger 24 is provided to cool the oil by passing the oil through a conduit cooled by the air flow.

The exchanger 24 is arranged in a channel 26, in which channel 26 a so-called third flow F3 circulates. The passage 26 may extend circumferentially over all or a portion of the turbine (i.e., over an angle of 360 ° or less about the axis 8). Similarly, the heat exchanger 24 may occupy all or a portion of the channel 26, extending over a large angular portion (in particular 360 °) around the axis 8.

The channel 26 has an opening 28 which opens into the space through which the secondary flow F2 passes. The opening 28 is defined by a fairing, upstream by an upstream edge 30, and downstream by a downstream edge 32. The upstream edge 30 and the downstream edge 32 are at substantially the same radial height to prevent any foreign elements present in the flow F from turning toward the exchanger 24.

Thus, the openings 28 are such that when the openings are formed upstream of the channel 26, the velocity of the flow F3 has a significant radial component. In particular, this ensures that the velocity of the flow F3 is much lower than the velocity of the main flow F1 in the flow channel 12 as the flow passes through the exchanger. Thus, the geometry may be such that when the Mach number of stream F1 is between 0.45 and 0.6 (typically 0.5), the Mach number of stream F3 observed by the exchanger is well below 0.3. These values ensure effective cooling at cruising speeds.

The opening 28 may be fitted with a protective grille (not shown) or a baffle to open or close the inlet of the secondary flow F2 to the channel 26.

The compressor 14 is constituted by an upstream portion 14.1 comprising blades arranged in the duct 12 and a downstream portion 14.2. The downstream portion 14.2 comprises at least one rotor blade 20, 22 which generates a flow F3, creating a vacuum at the opening 28.

Once the third stream has passed through the exchanger 24, the third stream F3 returns to the downstream compressor section 14.2. In the first embodiment, the flow F3 merges with the main flow F1.

The radial height H of the blades 20, 21, 22 in the downstream portion 14.2 is between 1.5 and 4 times the radial height H of the blades 16, 17, 18, 19 in the upstream portion 14.1.

Fig. 1 shows in dashed lines the possible positions of an annular row of vanes 34, which are fixed about the axis 8 and straighten the flow F2. The vanes 34 and propeller blades 4 may have a variable orientation (direction around the largest dimension of the vanes and propeller blades).

The number of compressor blades 14 forming the downstream portion 14.2 (i.e. the number of blades observed by the third flow F3 exiting the channel 26) may vary.

Thus, in the embodiment shown in fig. 2, the entire compressor is arranged in the downstream portion 14.2, no compressor blades being arranged in the flow path 12.

The flow channel 12 may be, for example, axially shorter than in fig. 1 and may comprise only one row of inlet vanes 11 and support arms 13.1.

In a third embodiment of the invention, the bypass channel 36 is arranged to divert a portion of the flow towards the secondary flow F2.

A bypass 36 is arranged downstream of the compressor 14 to divert the portion of the flow exiting the compressor towards the secondary flow F2. In this way, the secondary separation nozzle 38 separates the flow observed by the downstream portion 14.2 of the compressor into a main flow F1, which continues to the high-pressure compressor and the combustion chamber, and a fourth flow F4, which returns to the secondary flow F2.

The outer shroud 40 radially defines the passage 26 and the bypass 36, and internally defines the flow F2.

In addition to the heat exchange function of the fourth flow with the heat exchanger, the fourth flow F4 has a thrust function that supplements the flow F2. Any pressure loss in the flow F3 due to interaction with the exchanger 24 is compensated by the vanes in the downstream portion 14.2 to re-establish sufficient pressure in the fourth flow F4 before the fourth flow returns to the secondary flow F2.

Thus, a compromise is achieved between a low pressure in the conduit 26 sufficient to promote heat exchange and a high pressure in the bypass 36 sufficient to promote thrust.

Fig. 4 shows a variant in which a single row of blades occupies the downstream portion 14.2 of the compressor 14. To structurally support the outer shroud 40 without stator vanes in the downstream portion 14.2, structural arms 42 may be disposed in the bypass 36 and/or in the passage 26 (not shown).

Fig. 4 also shows an aspect of the embodiment shown in fig. 3, namely the isolation between the main flow F1 and the third flow F3. Thus, the fourth flow F4 of fig. 3 is the third flow F3: each air particle passing through the exchanger 24 continues to flow through its path through the bypass 36.

Means such as seals or rings are provided to isolate the main flow from the third flow.

In this regard, fig. 4 shows a circumferential ring 44 that circumferentially couples the blades of the rotor blade row of the downstream portion 14.2 close to each other. The ring is radially arranged at the nozzle 38 and the shroud 13.

In a variant, not shown, in which the downstream portion 14.2 comprises a plurality of rows of blades (as shown in fig. 3), rings of the same type may be arranged on all the rows of blades. Alternatively, the ring may be provided only on the last blade row of the downstream portion 14.2 in preparation for a flow separation between the main flow to the high pressure compressor and the bypass 36, thereby only partly isolating the main flow from the third flow.

It should be noted that the present invention is not limited to the examples shown in the drawings. In particular, the teachings of the present invention are also applicable to turbines having shrouded propellers.

Each feature of each illustrated example is applicable to other examples. In particular, the number of blades in the downstream portion, whether there are bypasses, the presence of a position of a propeller or straightening vane, the presence of a circumferential ring, etc. may be applied from one embodiment to another.

Claims

1. A turbine (2) of the unducted fan (4) type, said turbine comprising:

-a separating nozzle (10) dividing the air flow (F) into a main flow (F1) and a secondary flow (F2);

-a compressor (14) compressing the main flow (F1); and

-An air/oil heat exchanger (24);

Characterized in that the exchanger (24) is arranged in a channel (26) traversed by a third flow (F3), the third flow (F3) being taken from the secondary flow (F2) upstream of the exchanger (24), and the third flow encountering at least one annular row of rotor blades (20, 22) of the compressor (14) downstream of the exchanger (24).

2. Turbine (2) according to claim 1, wherein the compressor (14) comprises an upstream portion (14.1) having at least one annular row of blades (16-19) and a downstream portion (14.2) having at least one annular row of blades (20-22), and only the downstream portion (14.2) of the compressor (14) is crossed by air from the third flow (F3) coming from the exchanger (24).

3. Turbine (2) according to claim 2, wherein the radial height (H) of the blades (20-22) of the downstream portion (14.2) is much greater than the radial height of the blades (16-19) of the upstream portion (14.1), preferably the blades (20-22) of the downstream portion (14.2) of the compressor (14) are 1.5 to 4 times radially higher than the blades (16-19) of the upstream portion (14.1).

4. A turbine (2) according to any one of claims 1 to 3, wherein the downstream portion (14.2) of the compressor (14) comprises a single annular row of rotor blades (20), the air encountering the downstream portion of the compressor downstream of the exchanger (24).

5. Turbine (2) according to any one of claims 1 to 4, wherein the downstream portion (14.2) of the compressor (14) comprises one or two annular rows of stator blades (21), the air encountering the downstream portion of the compressor downstream of the exchanger (24).

6. Turbine (2) according to any one of claims 1,4 or 5, characterized in that the air flow (F3) leaving the exchanger (24) passes through all annular rows of blades (18-22) of the compressor (14).

7. Turbine (2) according to any one of claims 1 to 6, wherein an opening (28) allows a portion of the secondary flow (F2) to be sucked in to form the third flow (F3), the opening (28) being non-ingestible.

8. The turbine (2) according to any one of claims 1to 7, wherein the channel (26) and the exchanger (24) extend circumferentially over 360 ° around the axis (8) of the turbine (2).

9. Turbine (2) according to any one of claims 1 to 8, wherein a propeller (4) is arranged upstream of the nozzle (10).

10. Turbine (2) according to any one of claims 1 to 9, wherein the vanes (34) straightening the secondary flow (F2) axially overlap with the downstream portion (14.2) of the compressor (14).

11. Turbine (2) according to any one of claims 1 to 8 or 10, wherein two propellers having opposite rotational directions are arranged at an axial position downstream of the exchanger (24).

12. Turbine (2) according to any one of claims 1 to 11, characterized in that a bypass (36) is arranged downstream of the downstream portion (14.2) of the compressor (14) to divert the portion of the flow leaving the compressor (14) towards the secondary flow (F2).

13. Turbine (2) according to the preceding claim, wherein means are provided for isolating the main flow (F1) from the third flow (F3) which flows into the passage (26) through the downstream portion (14.2) of the compressor (14) and then into the bypass (36).

14. Turbine (2) according to the preceding claim, wherein at least the last annular row of blades (22) of the downstream portion (14.2) has an intermediate circumferential ring (44).

15. A method for cooling oil of a turbine (2) having an unducted fan (4), the method comprising:

-dividing the air flow (F) into a main flow (F1) and a secondary flow (F2) by means of a separation nozzle (10), the main flow (F1) being compressed by one or more compressors (14) of the turbine (2), the secondary flow (F2) being located outside the one or more compressors (14);

-forming a third flow (F3) through an air/oil exchanger (24) by at least one annular row of rotor blades (20, 22) of a compressor (14), said third flow (F3) being taken from said secondary flow (F2).

16. A method of operating a turbine according to any one of claims 1 to 14, the method comprising a propeller rotation step during which the mach number of the primary flow (F1) and the secondary flow (F2) is 0.5, the mach number of the third flow (F3) being well below 0.3.